EP3851642A1 - Dichtungsanordnung zwischen brennkammer und schaufel und verfahren zu deren herstellung - Google Patents

Dichtungsanordnung zwischen brennkammer und schaufel und verfahren zu deren herstellung Download PDF

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Publication number
EP3851642A1
EP3851642A1 EP21152146.3A EP21152146A EP3851642A1 EP 3851642 A1 EP3851642 A1 EP 3851642A1 EP 21152146 A EP21152146 A EP 21152146A EP 3851642 A1 EP3851642 A1 EP 3851642A1
Authority
EP
European Patent Office
Prior art keywords
seal
shell
vane stage
gas turbine
forward face
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Granted
Application number
EP21152146.3A
Other languages
English (en)
French (fr)
Other versions
EP3851642B1 (de
Inventor
Steven D. PORTER
Caroline Karanian
John T. Ols
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
RTX Corp
Original Assignee
Raytheon Technologies Corp
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Raytheon Technologies Corp filed Critical Raytheon Technologies Corp
Publication of EP3851642A1 publication Critical patent/EP3851642A1/de
Application granted granted Critical
Publication of EP3851642B1 publication Critical patent/EP3851642B1/de
Active legal-status Critical Current
Anticipated expiration legal-status Critical

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Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D9/00—Stators
    • F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
    • F01D9/023—Transition ducts between combustor cans and first stage of the turbine in gas-turbine engines; their cooling or sealings
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • F01D11/005—Sealing means between non relatively rotating elements
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D9/00—Stators
    • F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F16—ENGINEERING ELEMENTS AND UNITS; GENERAL MEASURES FOR PRODUCING AND MAINTAINING EFFECTIVE FUNCTIONING OF MACHINES OR INSTALLATIONS; THERMAL INSULATION IN GENERAL
    • F16J—PISTONS; CYLINDERS; SEALINGS
    • F16J15/00—Sealings
    • F16J15/16—Sealings between relatively-moving surfaces
    • F16J15/32—Sealings between relatively-moving surfaces with elastic sealings, e.g. O-rings
    • F16J15/3284—Sealings between relatively-moving surfaces with elastic sealings, e.g. O-rings characterised by their structure; Selection of materials
    • F16J15/3288—Filamentary structures, e.g. brush seals
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/42—Continuous combustion chambers using liquid or gaseous fuel characterised by the arrangement or form of the flame tubes or combustion chambers
    • F23R3/50—Combustion chambers comprising an annular flame tube within an annular casing
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/42—Continuous combustion chambers using liquid or gaseous fuel characterised by the arrangement or form of the flame tubes or combustion chambers
    • F23R3/60—Support structures; Attaching or mounting means
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2240/00—Components
    • F05D2240/55—Seals
    • F05D2240/56—Brush seals
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00—Function
    • F05D2260/84—Redundancy
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R2900/00—Special features of, or arrangements for continuous combustion chambers; Combustion processes therefor
    • F23R2900/00012—Details of sealing devices
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R2900/00—Special features of, or arrangements for continuous combustion chambers; Combustion processes therefor
    • F23R2900/00018—Manufacturing combustion chamber liners or subparts

Definitions

  • This disclosure relates generally to gas turbine engines, and more particularly to seal assemblies between combustors and turbine sections of gas turbine engines.
  • Combustors such as those used in gas turbine engines, may generally include radially spaced inner and outer shells which define a combustion chamber therebetween. High-energy gas flow generated in the combustion chamber may be directed into a turbine section of the gas turbine engine to effect rotation of one or more turbines. Gas turbine engines may include sealing arrangements between the combustor and the turbine section.
  • a gas turbine engine includes a turbine section including a first vane stage.
  • the gas turbine engine further includes a combustor disposed forward of the first vane stage.
  • the combustor includes a combustion chamber in fluid communication with the first vane stage.
  • the combustion chamber is radially defined between a first shell and a second shell.
  • the first shell includes a first seal assembly at an aft end of the first shell.
  • the first seal assembly includes a conformal seal forming a first seal between the first shell and a forward face of the first vane stage.
  • the second shell includes a second seal assembly at an aft end of the second shell.
  • the second seal assembly includes a brush seal forming a second seal between the second shell and the forward face of the first vane stage.
  • the brush seal includes a brush seal backing plate, retaining ring, and a plurality of bristles retained between the brush seal backing plate and the retaining ring.
  • the plurality of bristles extend axially between the second shell and the forward face of the first vane stage.
  • the second seal is a secondary seal and the second seal assembly further includes a hard seal.
  • the hard seal forms a primary seal between the second shell and the forward face of the first vane stage.
  • the hard seal includes a radially extending face of the second shell in contact with the forward face of the first vane stage.
  • the hard seal is disposed radially between the brush seal and the combustion chamber.
  • the conformal seal includes a wearing end portion extending axially aft from the conformal seal and configured to wear down during initial operation of the gas turbine engine to form the first seal with the forward face of the first vane stage.
  • the gas turbine engine includes at least one aft liner panel mounted to the second shell radially adjacent the second seal assembly.
  • the combustor includes a single mounting point between the combustor and a fixed casing structure of the gas turbine engine.
  • the single mounting point is disposed between the second shell and the fixed casing structure.
  • a method for providing a seal between a combustor and a turbine section of a gas turbine engine includes providing the combustor including a combustion chamber radially defined between a first shell and a second shell. The method further includes forming a first seal between an aft end of the first shell and a forward face of a first vane stage of the turbine section with a first seal assembly including a conformal seal. The method further includes forming a second seal between an aft end of the second shell and the forward face of the first vane stage of the turbine section with a second seal assembly including a brush seal.
  • the brush seal includes a brush seal backing plate, retaining ring, and a plurality of bristles retained between the brush seal backing plate and the retaining ring.
  • the plurality of bristles extends axially between the second shell and the forward face of the first vane stage.
  • the second seal is a secondary seal and the second seal assembly further includes a hard seal.
  • the method further includes forming a primary seal between the second shell and the forward face of the first vane stage with the hard seal.
  • the hard seal includes a radially extending face of the second shell in contact with the forward face of the first vane stage.
  • the hard seal is disposed radially between the brush seal and the combustion chamber.
  • the conformal seal includes a wearing end portion extending axially aft from the conformal seal.
  • the method further includes wearing down the wearing end portion during initial operation of the gas turbine engine to form the first seal with the forward face of the first vane stage.
  • the combustor further includes at least one aft liner panel mounted to the second shell radially adjacent the second seal assembly.
  • a gas turbine engine includes a turbine section including a first vane stage.
  • the gas turbine engine further includes a combustor disposed forward of the first vane stage.
  • the combustor includes a combustion chamber in fluid communication with the first vane stage.
  • the combustion chamber is defined between an inner shell and an outer shell.
  • the inner shell includes an inner diameter seal assembly at an aft end of the inner shell.
  • the inner diameter seal assembly includes a hard seal forming a first inner diameter seal between the inner shell and the forward face of the first vane stage, and a brush seal, disposed radially inward of the hard seal, forming a second inner diameter seal between the inner shell and the forward face of the first vane stage.
  • the outer shell includes an outer diameter seal assembly at an aft end of the outer shell.
  • the outer diameter seal assembly includes a conformal seal forming an outer diameter seal between the outer shell and the forward face of the first vane stage.
  • the brush seal includes a brush seal backing plate, retaining ring, and a plurality of bristles retained between the brush seal backing plate and the retaining ring and the plurality of bristles extends axially between the inner shell and the forward face of the first vane stage.
  • the gas turbine engine 10 is schematically illustrated.
  • the gas turbine engine 10 is disclosed herein as a two-spool turbofan engine that generally includes a fan section 12, a compressor section 14, a combustor section 16, and a turbine section 18.
  • the fan section 12 drives air along a bypass flowpath 20 while the compressor section 14 drives air along a core flowpath 22 for compression and communication into the combustor section 16 and then expansion through the turbine section 18.
  • a turbofan gas turbine engine in the disclosed non-limiting embodiments, it should be understood that the concepts described herein are not limited to use with turbofans as the teachings may be applied to other types of turbine engines including those with three-spool architectures.
  • the gas turbine engine 10 generally includes a low-pressure spool 24 and a high-pressure spool 26 mounted for rotation about a longitudinal centerline 28 of the gas turbine engine 10 relative to an engine static structure 30 via one or more bearing systems 32. It should be understood that various bearing systems 32 at various locations may alternatively or additionally be provided.
  • the low-pressure spool 24 generally includes a first shaft 34 that interconnects a fan 36, a low-pressure compressor 38, and a low-pressure turbine 40.
  • the first shaft 34 is connected to the fan 36 through a gear assembly of a fan drive gear system 42 to drive the fan 36 at a lower speed than the low-pressure spool 24.
  • the high-pressure spool 26 generally includes a second shaft 44 that interconnects a high-pressure compressor 46 and a high-pressure turbine 48. It is to be understood that "low pressure” and "high pressure” or variations thereof as used herein are relative terms indicating that the high pressure is greater than the low pressure.
  • An annular combustor 50 is disposed between the high-pressure compressor 46 and the high-pressure turbine 48 along the longitudinal centerline 28.
  • the first shaft 34 and the second shaft 44 are concentric and rotate via the one or more bearing systems 32 about the longitudinal centerline 28 which is collinear with respective longitudinal centerlines of the first and second shafts 34, 44.
  • Airflow along the core flowpath 22 is compressed by the low-pressure compressor 38, then the high-pressure compressor 46, mixed and burned with fuel in the combustor 50, and then expanded over the high-pressure turbine 48 and the low-pressure turbine 40.
  • the low-pressure turbine 40 and the high-pressure turbine 48 rotationally drive the low-pressure spool 24 and the high-pressure spool 26, respectively, in response to the expansion.
  • the combustor 50 includes an annular outer shell 52 and an annular inner shell 54 spaced radially inward of the outer shell 52, thus defining an annular combustion chamber 56 therebetween.
  • the outer shell 52 includes an axially forward end 58 and an axially aft end 60 as well as a first surface 62 facing the combustion chamber 56 and a second surface 64 opposite the first surface 62.
  • the inner shell 54 includes an axially forward end 66 and an axially aft end 68 as well as a first surface 70 facing the combustion chamber 56 and a second surface 72 opposite the first surface 70.
  • the combustor 50 may further include a plurality of liner panels 73 mounted to the respective first surfaces 62, 70 of one or both of the outer shell 52 and the inner shell 54. It should be understood that relative positional terms, such as “forward,” “aft,” “upper,” “lower,” “above,” “below,” and the like are relative to the normal operational attitude of the gas turbine engine 10 and should not be considered otherwise limiting.
  • the turbine section 18 of the gas turbine engine 10 includes a first vane stage 74 which may be a forward-most vane stage of the turbine section 18.
  • the combustor 50 is axially forward of the first vane stage 74 and the combustion chamber 56 is in fluid communication with the first vane stage 74.
  • the first vane stage 74 includes an upper platform 76 which defines an outer diameter forward face 78 of the first vane stage 74 which is adjacent the axially aft end 60 of the outer shell 52.
  • the first vane stage 74 further includes a lower platform 80 which defines an inner diameter forward face 82 of the first vane stage 74 which is adjacent the axially aft end 68 of the inner shell 54.
  • the first vane stage 74 includes a plurality of vanes 84 that extend between the upper platform 76 and the lower platform 80.
  • Each vane of the plurality of vanes 84 includes a leading edge 86 facing toward the combustion chamber 56. The leading edge 86 encounters the high-energy gas flow generated in the combustion chamber 56 and directs that gas flow further downstream in the turbine section 18.
  • the outer shell 52 is in sealed communication with the first vane stage 74 via an outer diameter seal assembly 88 disposed at the axially aft end 60 of the outer shell 52.
  • the outer diameter seal assembly 88 includes a conformal seal 90 forming an outer diameter seal between the outer shell 52 and the outer diameter forward face 78 of the first vane stage 74.
  • the outer shell 52 may include a rib 92 which extends radially outward from the second surface 64 of the outer shell 52 and which is axially spaced from the axially aft end 60 of the outer shell 52.
  • the conformal seal 90 may be disposed between the rib 92 and the outer diameter forward face 78 of the first vane stage 74 on the second surface 64 of the outer shell 52.
  • the conformal seal 90 may extend axially aft from the rib 92 over a radially extending gap 94 between the outer shell 52 and the first vane stage 74.
  • the conformal seal 90 may include a wearing end portion 96 extending axially aft from the conformal seal 90 and configured to wear down during initial operation of the gas turbine engine 10 to form the outer diameter seal with the outer diameter forward face 78 of the first vane stage 74.
  • the conformal seal 90 may include a plurality of cooling holes 98 which extend between a radially outer surface 100 and a radially inner surface 102 of the conformal seal 90.
  • the plurality of cooling holes 98 may be in fluid communication with the radially extending gap 94 so as to provide cooling air to the radially extending gap 94.
  • the inner shell 54 is in sealed communication with the first vane stage 74 via an inner diameter seal assembly 104 disposed at the axially aft end 68 of the inner shell 54.
  • the inner diameter seal assembly 104 may include a hard seal 106 forming an inner diameter seal between the inner shell 54 and the inner diameter forward face 82 of the first vane stage 74.
  • the hard seal 106 may be defined by a radially extending surface 108 of the inner shell 54 at the axially aft end 68 which is in contact with the inner diameter forward face 82 of the first vane stage 74.
  • the inner diameter seal assembly 104 includes a brush seal 110 forming an inner diameter seal between the inner shell 54 and the inner diameter forward face 82 of the first vane stage 74.
  • the inner diameter seal assembly 104 may include both the hard seal 106 forming a first inner diameter seal and the brush seal 110 forming a second inner diameter seal.
  • the inner diameter seal assembly 104 may include the brush seal 110 and may not include the hard seal 106.
  • the brush seal 110 may be disposed radially inward of the hard seal 106, however, the brush seal 110 is not limited to such a location and may be, for example, disposed radially outward of the hard seal 106.
  • the brush seal 110 may include a brush seal backing plate 112, a retaining ring 114, and a plurality of bristles 116 radially retained between the brush seal backing plate 112 and the retaining ring 114.
  • the inner shell 54 may include a rib 118 which extends radially inward from the second surface 72 of the inner shell 54 and which is axially spaced from the axially aft end 68 of the inner shell 54.
  • the brush seal 110 may be disposed between the rib 118 and the inner diameter forward face 82 of the first vane stage 74 on the second surface 72 of the inner shell 54.
  • the plurality of bristles 116 may extend axially between the inner shell 54 and the inner diameter forward face 82 of the first vane stage 74 so as to contact the inner diameter forward face 82 with a distal end 120 of the plurality of bristles 116.
  • the brush seal 110 may have improved contact with the inner diameter forward face 82 of the first vane stage 74, as compared to the hard seal 106 and, therefore, may provide greater sealing between the inner shell 54 and the first vane stage 74 relative to an inner diameter seal assembly including a hard seal alone.
  • outer diameter seal assembly 88 is discussed above as including the conformal seal 90 and the inner diameter seal assembly 104 is discussed above as including one or both of the hard seal 106 and the brush seal 110, it should be understood that other configurations of the seals 90, 106, 110 may be used.
  • the outer diameter seal assembly 88 may include one or both of the hard seal 106 and the brush seal 110 while the inner diameter seal assembly 104 may include the conformal seal 90.
  • At least one liner panel of the plurality of liner panels 73 may be mounted to the outer shell 52 radially inward (e.g., radially adjacent) of the outer diameter seal assembly 88.
  • at least one liner panel of the plurality of liner panels 73 may be mounted to the inner shell 54 radially outward (e.g., radially adjacent) of the inner diameter seal assembly 104.
  • the combustor 50 may include a single mounting point between the combustor 50 and a fixed casing structure 122 of the gas turbine engine 10.
  • the inner shell 54 may include an inner diameter leg 124 extending from the second surface 72 of the inner shell 54 and mounted to the fixed casing structure 122 via one or more fasteners 126.
  • the combustor 50 may experience some rocking and twisting relative to the first vane stage 74 during operation of the gas turbine engine 10.
  • Inclusion of the brush seal 110 in the inner diameter seal assembly 104 may accommodate the rocking and twisting while still provided adequate sealing between the inner shell 54 and the first vane stage 74.

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  • Engineering & Computer Science (AREA)
  • General Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP21152146.3A 2020-01-17 2021-01-18 Dichtungsanordnung zwischen brennkammer und schaufel und verfahren zu deren herstellung Active EP3851642B1 (de)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
US16/746,210 US20210222878A1 (en) 2020-01-17 2020-01-17 Combustor to vane sealing assembly and method of forming same

Publications (2)

Publication Number Publication Date
EP3851642A1 true EP3851642A1 (de) 2021-07-21
EP3851642B1 EP3851642B1 (de) 2024-03-06

Family

ID=74187193

Family Applications (1)

Application Number Title Priority Date Filing Date
EP21152146.3A Active EP3851642B1 (de) 2020-01-17 2021-01-18 Dichtungsanordnung zwischen brennkammer und schaufel und verfahren zu deren herstellung

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US (1) US20210222878A1 (de)
EP (1) EP3851642B1 (de)

Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US5480162A (en) * 1993-09-08 1996-01-02 United Technologies Corporation Axial load carrying brush seal
US5848874A (en) * 1997-05-13 1998-12-15 United Technologies Corporation Gas turbine stator vane assembly
US6357752B1 (en) * 1999-10-15 2002-03-19 General Electric Company Brush seal
EP3450692A1 (de) * 2017-08-30 2019-03-06 United Technologies Corporation Bugwellenkühlung mit konformer dichtung
EP3511622A1 (de) * 2018-01-16 2019-07-17 Rolls-Royce plc Brennkammeranordnung

Family Cites Families (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US8777563B2 (en) * 2011-01-31 2014-07-15 General Electric Company Axial brush seal
US10669939B2 (en) * 2016-10-26 2020-06-02 Raytheon Technologies Corporation Combustor seal for a gas turbine engine combustor

Patent Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US5480162A (en) * 1993-09-08 1996-01-02 United Technologies Corporation Axial load carrying brush seal
US5848874A (en) * 1997-05-13 1998-12-15 United Technologies Corporation Gas turbine stator vane assembly
US6357752B1 (en) * 1999-10-15 2002-03-19 General Electric Company Brush seal
EP3450692A1 (de) * 2017-08-30 2019-03-06 United Technologies Corporation Bugwellenkühlung mit konformer dichtung
EP3511622A1 (de) * 2018-01-16 2019-07-17 Rolls-Royce plc Brennkammeranordnung

Also Published As

Publication number Publication date
US20210222878A1 (en) 2021-07-22
EP3851642B1 (de) 2024-03-06

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